Gas guns generally have low rates of fire because of the long periods of time required to fill the combustion chamber with the necessary gases unless expensive and leakage-prone gas filled cartridges are used. Light gas guns in particular can provide very high muzzle velocities but cannot presently load propellants quickly enough to achieve the high rates of fire required by an effective weapon. The friction and compression heating caused by extremely rapid transfer of room temperature gases results in high temperatures that especially impede injection of the necessary quantity of propellants into relatively small volume, closed combustion chambers. Combustion driven presses for powdered metal parts and other production as well as very fast and reliable combustion driven switches have similar cycling limitations with present injection methods.
Liquid fuel rocket engines generally employ turbomachinery for pressurizing and/or gasifying the liquid propellants prior to injection into the main rocket nozzle. Furthermore, one or more of the propellant components may be adapted to cool the main rocket nozzle and heat the cryogenic propellant through associated plumbing circuitry. Liquid fuel and liquid oxidizer are provided from pressurized tanks at relatively low-pressure to separate sections within a rotor system driven by a relatively low-pressure ratio turbine that is powered by combustion effluent generated by a precombustor. A rocket engine frequently uses primary and secondary rotary injectors for injecting fuel and oxidizer propellant components into a first combustion chamber, and the effluent drives a turbine that rotates the rotary injectors. The mixture within the first combustion chamber is preferably fuel-rich so as to reduce the associated combustion temperature, and the fuel-rich effluent mixes in a second combustion chamber with additional oxidizer injected by a third rotary injector so as to generate a high temperature effluent suitable for propulsion. The rotary injectors are adapted to isolate the low-pressure propellant supply from the relatively high pressures in the respective combustion chambers. Rocket engines use steady state injection into open combustion chambers with much lower injection and combustion pressures than either gas gun or combustion driven applications.
A considerable amount of heat is transferred in all designs of rocket engines. The principle objective of high-temperature rocket design is to safely limit the heat transfer to the materials in critical hot sections such as the injector, combustion chamber, throat, and nozzle. The walls have to be cooled sufficiently to not exceed their safe allowable operating limit. Erosion, usually the result of combined oxidation and chemical interaction with the hot combustion gases, should not damage the walls, and the walls should be capable of withstanding the extreme thermal shock caused by the sudden onset of a high heat flux from combustion ignition. The materials comprising the thrust chamber devices must also be capable of resisting the thermal stresses induced by the heat transfer and thermal gradients.
Actively-cooled liquid propellant thrust chambers have provisions for cooling some or all of the components in contact with the hot combustion gases, such as the chamber walls, nozzle walls and injector faces. A cooling jacket or cooling coil often consists of separate inner and outer walls or a bundled assembly of continuous, contoured tubes. The inner wall confines the combustion gases, and the space between the inner and outer walls serves as the coolant passage. Regenerative cooling is a form of active cooling and is used for engines where one of the propellant constituents is circulated through cooling passages around the thrust chamber prior to injection and burning of the propellant in the combustion chamber. Regenerative cooling in bipropellant engines uses either the fuel or oxidizer as the cooling fluid. Therefore, the thermal energy absorbed by the coolant is not wasted as it augments the initial energy content of the propellant prior to injection, thereby increasing the exhaust velocity and propulsive performance. Radiation cooling is typically used in monopropellant thrust chambers, some gas generators and for nozzle exhaust sections. Radiation cooling is a simple, lightweight cooling method, which is commonly employed in low-temperature rocket engines, such as hydrazine (monopropellant) spacecraft maneuver and attitude control systems, where the maximum chamber temperature is only about 650 degrees C.
The fundamental principle that allows a hybrid rocket to burn is that in steady state operation, the fuel surface is constantly generating a melt layer, which in turn generates vapor as more heat is added or the heat causes the fuel to sublime directly to vapor from solid phase. The method of improving combustion of a hybrid rocket by gasifying liquid oxygen (LOX) as it enters the hybrid motor, before it contacts the hybrid fuel, is comprised of connecting at least one O2-driven hybrid heater to the motor such that its exhaust stream intersects and mixes directly with the LOX stream. Gaseous oxygen (GOX) is provided to the hybrid heater during the entire burn of the rocket. The hybrid heater is preferably ignited with electrical current.
Liquid injection of fuel into supersonic combustors has experienced difficulty in achieving vaporization of the liquid droplets, followed by gas mixing of this vapor with the surrounding air so that complete, molecular-scale mixing and combustion can take place inside the combustor. If the fuel droplets are small enough to vaporize quickly, they are carried along with the flow and because there is little relative velocity between the vaporizing droplet and the surrounding flow, there is no driving force for mixing the air with the vapor fuel except by molecular diffusion, which is very slow compared with fluid dynamic mixing. If the fuel droplets are large and a relative velocity can be maintained with respect to the surrounding air to promote mixing, a large amount of heat is required to vaporize the droplets and fuel vapor is formed at a relatively slow rate compared with the same mass of fuel dispersed in smaller droplets. Thus for either large or small liquid droplets of fuel the final mixing of fuel and air on a molecular scale necessary for combustion is a slow process compared with the approximately 10 ms time scale to flow through a supersonic combustor.
A high-pressure pump and delivery system provides a method of utilizing both pumped LNG and compressed NG in a Diesel type fuel injection system. As the truck's engine requires fuel, the LNG is vaporized and supplied to the engine at a pre-determined pressure, with the desired pressure being a function of the engine's specific design. These engines are generally designed to operate at pressures between 200 psig and 2,000 psig with potential for as much as 3000 psig.
There is a need in the art for an improved, high rate transient closed-system cryogenic injection system.